This disclosure generally relates to a gas turbine engine, and more particularly to rotor blades that improve gas turbine engine performance.
Gas turbine engines, such as turbofan gas turbine engines, typically include a fan section, a compressor section, a combustor section and a turbine section. During operation, air is pressurized in the compressor section and mixed with fuel in the combustor section for generating hot combustion gases. The hot combustion gases flow through the turbine section which extracts energy from the hot combustion gases to power the compressor section and drive the fan section.
Many gas turbine engines include axial-flow type compressor sections in which the flow of compressed air is parallel to the engine centerline axis. Axial-flow compressors utilize multiple stages to obtain the pressure levels needed to achieve desired thermodynamic cycle goals. A typical compressor stage consists of a row of moving airfoils (called rotor blades) and a row of stationary airfoils (called stator vanes). The flow path of the axial-flow compressor section decreases in cross-sectional area in the direction of flow to reduce the volume of air as compression progresses through the compressor section. That is, each subsequent stage of the axial flow compressor decreases in size to maximize the performance of the compressor section.
One design feature of an axial-flow compressor section that may affect compressor performance is tip clearance flow. A small gap extends between the tip of each rotor blade and a surrounding shroud in each compressor stage. Tip clearance flow is defined as the amount of airflow that escapes between the tip of the rotor blade and the adjacent shroud. Tip clearance flow reduces the ability of the compressor section to sustain pressure rise and may have a negative impact on stall margin (i.e., the point at which the compressor section can no longer sustain an increase in pressure such that the gas turbine engine stalls).
Airflow escaping through the gaps between the rotor blades and the shroud can create gas turbine engine performance losses. In the middle and rear stages of the compressor section, blade performance and operability of the gas turbine engine are highly sensitive to the lower spans (i.e., decreased size) of the rotor blades and the corresponding high clearance to span ratios. Disadvantageously, prior rotor blade airfoil designs have not adequately alleviated the negative effects caused by tip clearance flow.